Liquid rocket engine tap-off power source

ABSTRACT

A liquid rocket engine integrates tap-off openings at a combustion chamber wall to direct exhaust from the combustion chamber to a tap-off manifold that provides the exhaust to one or more auxiliary systems, such as a turbopump that pumps oxygen and/or fuel into the combustion chamber. The tap-off opening passes through a fuel channel formed in that combustion chamber exterior wall and receives fuel through a fuel opening that interfaces the fuel channel and tap-off opening. The tap-off manifold nests within a fuel manifold for thermal management. The fuel channel directs fuel into the combustion chamber through fuel port openings formed in the combustion chamber, the fuel port openings located closer to a headend of the combustion chamber than the tap-off openings.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.17/324,000, filed May 18, 2021, entitled “Liquid Rocket Engine Tap-OffPower Source,” naming Thomas Edward Markusic and Anatoli AlimpievichBorissov as inventors, which application is a continuation of Ser. No.16/584,449, filed Sep. 26, 2019, now U.S. Pat. No. 11,008,977, issuedMay 18, 2021, entitled “Liquid Rocket Engine Tap-Off Power Source,”naming Thomas Edward Markusic and Anatoli Alimpievich Borissov asinventors, which applications are incorporated herein by reference inits entirety.

BACKGROUND OF THE INVENTION Field of the Invention

The present invention relates in general to the field of liquid rocketengines, and more particularly to a liquid rocket engine with tap-offpower source.

Description of the Related Art

Liquid rocket engines generate thrust by burning propellant, such asfuel mixed with liquid oxygen, in a combustion chamber at extremely hightemperatures and pressures, and exhausting the combustion gases througha throat and out a nozzle to produce a supersonic airflow. Generally,rockets launch payloads, such as satellites, into orbit by generatingthrust in excess of the weight of the rocket, fuel and oxygen storedonboard the rocket and the payload. By minimizing rocket structuralweight and maximizing efficient use of propellant, payload weightimproves for a desired orbit. Firefly Aerospace Inc. has invented across-impinged propellant injection and a cooling channel arrangementthat improve payload, as described in the following two patentapplications, which are incorporated herein as though fully set forth:“Liquid Rocket Engine Cooling Channels,” U.S. patent application Ser.No. 16/256,210, and “Liquid Rocket Engine Cross Impinged PropellantInjection,” U.S. patent application Ser. No. 16/256,206, both by AnatoliAlimpievich Borissov and Thomas Edward Markusic, the inventors hereof.

To provide fuel and oxygen flow to the combustion chamber with adequatepressure, liquid rocket engines typically have a turbopump that feedsfuel to a fuel manifold and oxygen to an oxygen manifold. As describedabove, the fuel generally passes through cooling channels formed in theliquid rocket engine thruster body to help manage thruster body thermalconstraints through regenerative absorption of excess thermal energy.The turbopump is typically powered by a small liquid rocket engine, alsoknown as a gas-generator, which exhausts into the turbopump turbineblades to generate power that pressurizes fuel and oxygen. The use of aseparate small liquid rocket engine to power the turbopump introduces anumber of difficulties and complexities. For example, the separateliquid rocket engine adds weight to the rocket and consumes fuel thatmight otherwise lift the payload. The added weight typically includes asystem of hydraulic valves used to manage propellants consumed by thesmall liquid rocket engine. The hydraulic valves also increase systemcost and complexity while offering many additional potential failureinstances, such as where turbopump pressure drops resulting ininsufficient cooling of the thruster body and catastrophic failure ofthe rocket. Another difficulty that arises with use of a small liquidrocket engine is that the exhaust that feeds into the turbopump tends tohave a fairly high level of soot, which impacts turbopump efficiency andlifespan.

Another option available to power the turbopump is to tap combustiongases at the combustion chamber to feed to the turbopump. Such so called“tap-off” systems were researched and tested in the 1960 era withminimal success. Although these early tap-off systems avoided theexpense and complexity of hydraulic valves to manage a gas generator,the difficulties associated with routing combustion gases from thecombustion chamber to the turbopump proved too substantial to permit aflight version. For example, tap-off gas conditions, such as pressure,temperature and gas composition, have to be managed within operatingconstraints of the turbopump throughout the flight envelope. The moreexpensive and complex gas generators became the only practicalalternative for commercial rocket flight.

SUMMARY OF THE INVENTION

Therefore, a need has arisen for a system and method which tapscombustion chamber energy to power liquid rocket engine auxiliarysystems.

In accordance with the present invention, a new system and method areprovided which substantially reduce the disadvantages and problemsassociated with previous methods and systems that tap combustion chamberenergy to power liquid rocket engine auxiliary systems, such asturbopump that provides fuel and oxygen to the liquid rocket engineunder pressure. Tap-off openings, formed in a combustion chamber wall,feed combustion gases to a tap-off manifold nested within a fuelmanifold and interfaced as a power source with auxiliary systems, suchas a turbopump. The tap-off openings pass through a cooling fin that isdisposed at the combustion chamber exterior in a fuel channel and thatincludes an opening from the fuel channel to the tap-off opening forfuel injection into the tap-off combustion gases. Both a toroidal vortexformed in the combustion chamber by oxygen and fuel injection and arelative location of fuel injection port openings to the tap-offopenings provide an oxygen rich tap-off flow to efficiently burn fuelinjected at the tap-off opening within the tap-off manifold.

More specifically, a liquid rocket engine thruster body generates thrustby burning fuel and oxygen injected into a combustion chamber to createa supersonic flow out a throat and nozzle from the combustion chamber. Aturbopump pumps fuel and liquid oxygen under high pressure into thecombustion chamber with power provided by a tap-off manifold thatreceives tap-off gases from the thruster body combustion chamber.Tap-off gases pass from the combustion chamber to the tap-off manifoldthrough tap-off openings cooled by fuel in fuel channels integrated inthe thruster body. For instance, tap-off openings are drilled through acooling fin formed in the fuel channel and interface with fuel disposedin the fuel channel through a tap-off fuel injection opening that addsfuel to the tap-off gases as the tap-off gases proceed through thetap-off opening. A cross-fire propellant injection system of thecombustion chamber generates an oxygen rich tap-off gas flow for tap-offopenings located below fuel injection openings relative to thecombustion chamber headend. The oxygen rich tap-off gases mix withinjected fuel and burn in a flameless mode that generates an optimalsoot-free exhaust to feed the turbopump. The tap-off manifold nestswithin a fuel manifold to maintain the tap-off manifold within thermalconstraints by transferring excess thermal energy to fuel in the fuelmanifold. Fuel guides formed in the outer surface of the tap-offmanifold enhance thermal exchange and guide fuel to fuel channels formedin the thruster body.

The present invention provides a number of important technicaladvantages. One example of an important technical advantage is that aliquid rocket engine has auxiliary systems, such as a turbopump thatpressurizes fuel and oxygen, powered by combustion chamber gases insteadof a separate gas generator. Tap-off power supplied to auxiliary systemsreduces rocket structural weight by eliminating the separate gasgenerator and removes a layer of complexity by eliminating hydraulicvalves associated with the separate gas generator. As a result,manufacture costs are reduced and liquid rocket engine reliability isincreased. By locating tap-off openings proximate to fuel port openings,combustion chamber gases enter the tap-off openings with a heated andoxygen-rich mixture that cleanly burns additional injected fuel in avolume distributed (flameless) mode that produces a better gas feed forthe turbopump with reduced soot content. Metallic structures of thetap-off opening and manifold are protected from thermal damage withregenerative thermal transfer to fuel of the fuel channels and the fuelmanifold in which the tap-off manifold is nested.

BRIEF DESCRIPTION OF THE DRAWINGS

The present invention may be better understood, and its numerousobjects, features and advantages made apparent to those skilled in theart by referencing the accompanying drawings. The use of the samereference number throughout the several figures designates a like orsimilar element.

FIG. 1 depicts a side view of a liquid rocket engine configured to powera turbopump with heated gases provided from the liquid rocket engine'scombustion chamber through a tap-off supply;

FIG. 2 depicts a side cutaway view of the liquid rocket engineconfigured to tap-off combustion chamber gases formed with a toroidalvortex flow of cross impinged fuel port openings in a side combustionchamber wall and an oxygen injector centrally located at the combustionchamber headend;

FIG. 3 depicts a side perspective exploded view of alignment of a nestedmanifold assembly at the headend of the thruster body;

FIG. 4 depicts a side view of an example arrangement of fuel and oxygenfeeds into a nested manifold assembly and tap-off supply out of thenested manifold assembly;

FIG. 5 depicts an upper side perspective view of the thruster body withthe nested manifold removed;

FIG. 6 depicts a side cutaway view of the thruster body that illustratesthe relationship between oxygen, fuel and tap-off openings within thecombustion chamber;

FIG. 7 depicts a side cutaway view of a nested manifold assembly with afuel manifold encircling a tap-off manifold to provide thermalmanagement by transfer of excess thermal energy from the tap-offmanifold to fuel disposed within the fuel manifold cavity;

FIG. 8 depicts an upper side perspective view of a tap-off and fuelmanifold inner portion before assembly into the nested manifoldassembly;

FIG. 9 depicts a side perspective view of thruster body fuel channelshaving the outer surface layer removed;

FIG. 10 depicts a close-up cut away view of fuel guides formed in theexternal surface of a combustion chamber that illustrates one exampleembodiment having tap-off openings drilled through tap-off fins; and

FIG. 11 depicts an example of a toroidal vortex formed by crossimpingement of injected fuel and oxygen in a combustion chamber relativeto tap-off openings.

DETAILED DESCRIPTION

A liquid rocket engine generates thrust with a more simple and reliabletap-off auxiliary power supply. Heated gases generated in a combustionchamber are tapped through tap-off openings formed in the combustionchamber wall and redirected through a tap-off manifold to a turbopumpthat applies the heated gases to drive fuel and oxygen pumps that feedthe combustion chamber. Tap-off thermal management is provided through ahighly developed external heat transfer surface by nesting the tap-offmanifold within the fuel manifold so that the tap-off manifold isregeneratively cooled by fuel flowing from the fuel source towardsthruster body fuel channels. Within a combustion chamber having fuelinjected from fuel port openings formed at cooling channels integratedin a side wall to impinge against oxygen injected at a headend towardsthe side wall, a toroidal vortex forms that allows an oxygen rich gasmixture to be drawn where the fuel port openings are disposed closer tothe headend than the tap-off openings. Fuel injection from the coolingchannels into the tap-off opening evaporates to further cool the tap-offgases and then mixes with the oxygen-rich tap-off gases through axialflow encouraged by the tap-off manifold for volume distributed(flameless) burning with the resulting gas having a near-idealcomposition for turbopump intake with minimal soot. Reliable andefficient auxiliary power supply is provided to the turbopump with thetap-off manifold safely sealed at the combustion chamber to preventleakage under high working fuel pressure.

Referring now to FIG. 1 , a side view depicts a liquid rocket engine 10configured to power a turbopump 14 with heated gases provided from theliquid rocket engine's combustion chamber through a tap-off supply 12.Liquid rocket engine 10 generates thrust by forcing oxygen and fuel withturbopump 14 from a liquid oxygen supply 18 and a fuel supply 20 into acombustion chamber for burning. For instance, liquid rocket engine 10 isbuilt into a rocket that houses a liquid oxygen tank interfaced withliquid oxygen supply 18 and a fuel tank interfaced with fuel supply 20.Turbopump 14 pumps oxygen provided from liquid oxygen supply 18 to aliquid oxygen feed 22 that interfaces with an oxygen manifold, and pumpsfuel provided from fuel supply 20 to a fuel feed 24 that interfaces witha fuel manifold. Fuel and oxygen pumped into liquid rocket engine 10burn to generate combustion gases having high temperatures and pressuresthat are forced through a throat and out a nozzle 26, which converts theinternal energy of the combustion gases to a kinetic energy ofsupersonic flow. Turbopump 14 exhausts tap-off gases at a tap-offexhaust 16 that feeds into nozzle 26 at the sonic flow. The energy ofthis flow translates to thrust applied by liquid rocket engine 10 to therocket in which it is coupled.

Referring now to FIG. 2 , a side cutaway view depicts liquid rocketengine 10 configured to tap-off combustion chamber 28 gases formed witha toroidal vortex flow of cross impinged fuel port openings 46 in a sidecombustion chamber wall 40 and an oxygen injector 34 centrally locatedat the combustion chamber 28 headend 36. In the example embodiment,thruster body 38 forms a combustion chamber 28 terminated at a headend36 to force high temperature and pressure exhaust gases through a throat30 and out a nozzle 26, thus generating a supersonic exhaust flow. Anoxygen manifold 32 located above headend 36 of thruster body 38 receivesliquid oxygen from turbopump 14 at high pressure to force the liquidoxygen into oxygen injector 34 for injection through oxygen portopenings 50 into combustion chamber 28. Similarly, a fuel manifold 42located exterior to combustion chamber wall 40 below headend 36 ofthruster body 38 receives fuel, such as kerosene, from turbopump 14 athigh pressure to force the fuel through fuel port openings 46 formedthrough combustion chamber wall 40. As is set forth below and in theincorporated patent applications in greater detail, fuel provided tofuel manifold 44 is forced through fuel channels integrated withincombustion chamber wall 40 to provide regenerative cooling to thrusterbody 38 before injection into combustion chamber 28. Tap-off openings 48formed in combustion chamber wall 40 provide a pathway for heatedcombustion gases from combustion chamber 28 into tap-off manifold 44from which the combustion gases exit to tap-off supply 12 as a powersource for turbopump 14. In the example embodiment, fuel port openings46 are symmetrically disposed about the inner surface circumference ofcombustion chamber wall 40 above symmetrically disposed tap-off openings48 relative to headend 36. As is set forth below, the toroidal vortexgenerated by impingement of fuel and oxygen injection within combustionchamber 28 results in an oxygen rich gas content at tap-off openings 48.

Referring now to FIG. 3 , a side perspective exploded view depictsalignment of a nested manifold assembly 52 at headend 36 of thrusterbody 38. Nested manifold assembly 52 has a central opening that fitsaround oxygen manifold 32 and couples to headend 36 with couplers 54,such as bolts that engage threads formed in headend 36. Nested manifoldassembly 52 fits around the upper circumference of thruster body 38 toexpose tap-off openings 48 drilled through into combustion chamber 28 toa cavity formed within tap-off manifold 44 that integrates in nestedmanifold assembly 52. At the upper circumference of thruster body 38 atheadend 36, fuel guides 56 are exposed that guide fuel into fuelchannels integrated in thruster body 38. Fuel manifold 42 integrated innested manifold assembly 52 is exposed to fuel guides 56 at the innercircumference of nested manifold assembly 52 above tap-off manifold 44so that fuel from fuel manifold 42 flows through fuel guides 56 and intofuel channels integrated in thruster body 38. Nested manifold assembly52 not only ensures a tight seal of tap-off manifold 44 but providescooling around the entire circumference of tap-off manifold 44 byflowing fuel into a fuel manifold cavity below tap-off manifold 44 andover top of tap-off manifold 44 into fuel guides 56.

Referring now to FIG. 4 , a side view depicts an example arrangement offuel and oxygen feeds into nested manifold assembly 52 and tap-offsupply out of nested manifold assembly 52. Nested manifold assembly 52fits over headend 36 at the upper portions of combustion chamber 28 fora simple manufacture process that provides a tight seal between fuel,oxygen and tap-off gases. Oxygen feed 22 provides oxygen to oxygenmanifold 32 at a top central location of combustion chamber 28 withclear separation from fuel feed 24 at a side of nested manifold assembly52. Fuel enters fuel feed 24 into a cavity of a fuel manifold withinnested manifold assembly 52 below a tap-off manifold within nestedmanifold assembly 52 so that the tap-off manifold is surrounded by fuelfor thermal exchange. Tap-off supply 12 provides a feed of heatedtap-off gases to feed turbopump 14.

Referring now to FIG. 5 , an upper side perspective view depictsthruster body 38 with nested manifold assembly 52 removed. The upperview illustrates how the example embodiment readily assembles yetmaintains separation of oxygen, fuel and tap-off gases during operation.Headend 36 defines a barrier at coupling of oxygen manifold 32 thatseparates cooling fuel flow into fuel guides 56 from the interior ofcombustion chamber 28. Nested manifold assembly 52 assembles overheadend 36 so that fuel guides 56 align with fuel guides within the fuelmanifold to direct fuel into fuel channels integrated in combustionchamber 28. Tap-off openings 48 communicate a tap-off manifold cavitywith combustion chamber 28 down and away from fuel guides 56 to provideseparation between fuel feeding into fuel channels and tap-off gasesexiting combustion chamber 28 through tap-off openings 48.

Referring now to FIG. 6 , a side cutaway view of thruster body 38illustrates the relationship between oxygen, fuel and tap-off openingswithin combustion chamber 28. As is explained in depth in theincorporated patent applications, oxygen port openings 50 and fuel portopenings 46 align to generate fuel and oxygen flows that impinge againsteach other, thus creating a toroidal vortex that mixes oxygen and fuelfor more efficient burning away from thruster body 38 inner surfaces.The top row of oxygen port openings 50 proximate the headend provideabout thirty percent (30%) of oxygen to create a strong near wall jet ofoxygen. The jet provides cooling of the headend and lateral walls of thecombustion chamber and help to form the vortex flow. The oxygenconcentration is enriched on the outer boundary of the vortex (near thewall of the combustion chamber) due to the elevated amount of oxygeninjected from the top set of oxygen ports. A cutaway of combustionchamber 28 inner surface above fuel port openings 46 illustrates fuelchannels 62 integrated in thruster body 38 that accept fuel from fuelguides 56 to route the fuel through thruster body 38 for regenerativecooling. Tap-off openings 48 are located below fuel port openings 46relative to headend 36 so that the toroidal vortex flow withincombustion chamber 28 will induce a heated and oxygen-rich combustionchamber gas content into tap-off openings 48. Each tap-off opening 48feeds combustion chamber gases through thruster body 38 and into atap-off manifold cavity 60 defined within tap-off manifold 44. As is setforth in greater detail below, tap-off opening 48 is formed through acooling fin structure within fuel channels 62 to provide cooling to themetal that defines the tap-off openings 48. Tap-off openings 48 areessentially drilled completely through combustion chamber 28 from theinterior surface to the exterior surface where tap-off manifold cavity60 accepts the heated gases. In contrast, fuel port openings 46 provideflow of fuel from fuel channels 62 into combustion chamber 28 and arethus drilled from only the interior surface of combustion chamber 28into the fuel channel 62.

In the example embodiment of FIG. 6 , fuel manifold 42 defines a fuelmanifold cavity 58 around tap-off manifold 44 that provides thermaltransfer from tap-off manifold 44 into fuel as the fuel enters fuelmanifold cavity 58 at its lower portion and proceeds over top of tap-offmanifold 44 to enter into fuel channels 62. Tap-off manifold cavity 60provides an axial flow of tap-off gases around the outer circumferenceof thruster body 38 to mix and burn oxygen and fuel as the tap-off gasesproceed towards tap-off supply 12. In particular, since the oxygen richcomposition provided from combustion chamber 28 through tap-off openings48 to tap-off manifold 44 is heated (as opposed to the relatively coldflow of oxygen provided to a gas generator), a more gradual andefficient nearly flameless burning of additional fuel added to tap-offmanifold 44 provides reduced soot in the exhaust that proceeds outtap-off supply 12 to turbopump 14.

Referring now to FIG. 7 , a side cutaway view depicts nested manifoldassembly 52 with fuel manifold 42 encircling tap-off manifold 44 toprovide thermal management by transfer of excess thermal energy fromtap-off manifold 44 to fuel disposed within fuel manifold cavity 58.Fuel within fuel manifold cavity 58 is pumped under pressure through afuel path 64 and into fuel channels integrated in thruster body 38. Thefuel channels 62 pass proximate the interior circumference of tap-offmanifold 44 so that thermal management by transfer of thermal energy toflowing fuel is provided about the entire exterior surface of tap-offmanifold 60.

Referring now to FIG. 8 , an upper side perspective view depicts atap-off and fuel manifold inner portion before assembly into a nestedmanifold assembly 52. Fuel guides 56 are machined into the outer surfaceof tap-off manifold cavity 60 to direct fuel from fuel manifold cavity58 along a fuel path 64 to fuel guides of the thruster body that feedfuel channels 62. Machining tap-off manifold cavity 60 and fuel manifoldcavity 58 on opposing sides of a contiguous metal piece providesstrength of the structure and reduced risk of leakages.

Referring now to FIG. 9 , a side perspective view depicts thruster body38 fuel channels 62 having the outer surface layer removed. Theincorporated patents describe manufacture of thruster body 38 bymachining fuel channels 62, filling the fuel channels 62 with wax andthen depositing nickel over top to define enclosed fuel channels. Thefuel channels 62 formed in thruster body 38 may be manufactured in amanner similar to that described by the incorporated patentapplications. In addition to drilling fuel port openings 46 to provide apathway for fuel injection from fuel channels 62 into combustion chamber28, tap-off openings 48 are drilled to provide access of combustiongases to a tap-off manifold coupled external to thruster body 38. Thus,one difference in the manufacture of tap-off openings 48 from fuel portopenings 46 is that the tap-off openings 48 pass completely throughthruster body 38, whereas fuel port openings 46 only provide a pathwayfrom fuel channels 62 into combustion chamber 28. To achieve anappropriate pathway for tap-off gas flow, a fuel guide structure 66 isincluded in fuel channels 62 that provide fuel to combustion chamber 28.Fuel flows from headend 36 of combustion chamber 28 through fuelchannels 62 to the end of nozzle 26 and then return to where fuel guidestructures 66 are formed for injection into combustion chamber 28through fuel port openings 46.

Near fuel guide structure 66, a tap-off fin 68 is machined between twoadjacent fuel channels 62 that bring fuel to fuel port openings 46.Tap-off openings 46 are drilled through the tap-off fin 68, which has aheight that extends to where the outer surface of thruster body 38 isformed by nickel deposition. During plating, wax fills the tap-offopening 48 so that a through opening is left from the interior surfaceof combustion chamber 28 to the outer surface of combustion chamber 28.As is described in greater detail below, fuel guide structure 66 definesa flow path of fuel from adjacent fuel channels 62 around tap-off fin 68for improved thermal transfer where combustion chamber 28 gases passthrough tap-off opening 48.

Referring now to FIG. 10 a close-up cut away view of fuel guides 56formed in the external surface of combustion chamber 28 illustrates oneexample embodiment having tap-off openings 48 drilled through tap-offfins 68. In the example embodiment, fuel guide structure 66 has atriangular shape that merges multiple fuel source paths 70 into twosingle paths traveling past where tap-off fin 68 is disposed withinadjacent fuel injection paths 72. During operation, fuel enters intofuel source paths 70 at combustion chamber 28 headend 36 so that allfuel channels 62 from headend 36 to fuel guide structure 66 are carryingfuel in the same direction, i.e., towards nozzle 26. After fuel guidestructure 66 forces merger of the fuel source paths 70, these pathsproceed to nozzle 26 where their direction is reversed to become fuelinjection paths 72. Once fuel travels along fuel injection path 72 tohit fuel guide structure 66, the fuel is forced out of fuel portopenings 46 and injected into combustion chamber 28.

In the example embodiment, tap-off fin 68 is machined from thruster body38 to create a structure that will support drilling of a tap-off opening48 through it and also have surface area exposed to fuel injection path72 that allows transfer of thermal energy from tap-off gases passingthrough tap-off opening 48 to fuel in fuel injection path 72. Inaddition, a tap-off opening fuel injection port 74 having an ellipticalor oval shape provides a path for fuel in fuel injection path 72 to passinto tap-off opening 48. Injecting fuel into tap-off opening 48 and/orthe tap-off manifold cavity helps to reduce tap-off gas temperature asthe fuel evaporates and also adds fuel to the oxygen rich gas thatenters tap-off opening 48 from combustion chamber 28. In variousembodiments, fuel injects directly into the tap-off opening from anelliptical shaped opening at the fuel channel with an expandingcircumference to aid atomization of the fuel. In one embodiment, atleast some of the fuel injects towards a center location of the tap-offmanifold to aid in mixing of the fuel with oxygen by the axial flow oftap-off gases around the tap-off manifold towards the tap-off source. Inalternative embodiments, alternative configurations for fuel injectioninto tap-off opening 48 may be used. For example, tap-off openings 48might be formed in a fuel source path 70 or have fuel injected from afuel source path 70 to provide a greater thermal gradient. Similarly,the location of tap-off opening 48 might vary relative to fuel portopenings 46 to achieve different fuel and oxygen mixture ratios. In theexample embodiment, the use of an elliptical or oval shaped tap-offopening fuel injection port 74 aids in maintaining structural integrityat tap-off fin 68.

Referring now to FIG. 11 , an example depicts a toroidal vortex 82formed by cross impingement of injected fuel flow 80 and oxygen 78 incombustion chamber 28 relative to tap-off openings 48. In the exampleembodiment, tap-off openings 48 are formed below fuel port openings 46relative to headend 36 of combustion chamber 28. As is explained ingreater detail in the incorporated patent applications, fuel flow 80injected from fuel ports 46 towards a central location of combustionchamber 28 impinges oxygen flow 78 injected from oxygen port openings 50to generate toroidal vortex 82, which mixes fuel and oxygen to burn awayfrom combustion chamber 28 walls. In part, toroidal vortex 82 isencouraged by oxygen port openings 50 aligned to inject oxygen parallelto headend 36 so that rotation of oxygen flow 78 promotes heating ofoxygen and mixing with fuel. As a result of the flow generated withincombustion chamber 28, oxygen to fuel ratio of combustion gas thatpasses by tap-off opening 48 is oxygen rich. In particular, placement oftap-off opening 48 below but in proximity to fuel port openings 46promotes a tap-off gas with a relatively high oxygen content. In variousembodiments, the relative placement of tap-off opening 48 may beadjusted so that oxygen content is of a desired amount in thecomposition of gas that enters tap-off opening 48.

Although the present invention has been described in detail, it shouldbe understood that various changes, substitutions and alterations can bemade hereto without departing from the spirit and scope of the inventionas defined by the appended claims.

What is claimed is:
 1. A method for powering an auxiliary system of aliquid rocket engine, the method comprising: pressurizing fuel in a fuelmanifold to pump the fuel through fuel channels formed in a wall of acombustion chamber and out fuel ports into the combustion chamber;injecting oxygen towards the combustion chamber wall from an oxygeninjector disposed at a headend of the combustion chamber; passingexhaust gas generated by burning fuel from the combustion chamber to atap-off manifold through tap-off openings formed in the combustionchamber wall; locating the fuel ports between the headend and thetap-off openings; cooling the tap-off manifold with fuel disposed aroundthe tap-off manifold; directing the exhaust gas from the tap-offmanifold to the auxiliary system; and powering the auxiliary system withthe exhaust gas.
 2. The method of claim 1 further comprising: directingat least some of the fuel from the fuel channels into the tap-offopenings.
 3. The method of claim 2 wherein the directing at least someof the fuel further comprises interfacing the fuel channel and tap-offopening with an elliptical-shaped opening formed between the fuelchannel and the tap-off opening.
 4. The method of claim 1 wherein theauxiliary system comprises a turbopump, the method further comprising:pumping the fuel into the fuel manifold with the turbopump; pumping theoxygen to the oxygen injector with the turbopump; and directing exhaustof the turbopump to an exhaust from the combustion chamber.
 5. Themethod of claim 1 wherein: at least some of the oxygen injects along theheadend substantially perpendicular to the combustion chamber wall; andexhaust entering the tap-off openings has an oxygen-rich composition. 6.The method of claim 5 wherein fuel directed into the tap-off openingmixes with the exhaust to establish a substantially volume distributedcombustion.